Gas turbine engine with bypass TOBI cooling/purge flow and method

Abstract

A gas turbine engine is provided that includes a high pressure compressor (HPC), a combustor section, a high pressure turbine (HPT), and a bypass tangential on board injector (TOBI) system. The combustor section has a combustor. A core gas path extends through the HPC, the combustor section, and the HPT. The bypass TOBI system extends circumferentially around the engine axial centerline, and has a plurality of nozzles, inner and outer radial sides, a plurality of first type and second type radial passages configured to allow the gas from the HPC to pass from the inner radial side of the bypass TOBI system to the outer radial side of the bypass TOBI system, wherein the first type radial passages are differently configured from the second type radial passages.

The invention claimed is: 
     
       1. A gas turbine engine having an axial centerline, comprising:
 a high-pressure compressor (HPC) having an HPC rotor stage and an HPC stator vane stage, the HPC rotor stage having a disk and a plurality of HPC rotor blades, and the HPC stator vane stage having a plurality of HPC stator vanes, the HPC having an HPC discharge at an aft end of the HPC; 
 a combustor section having a combustor; 
 a high-pressure turbine (HPT) having an HPT rotor stage and an HPT stator vane stage, the HPT rotor stage having a disk and a plurality of HPT rotor blades, and the HPT stator vane stage having a plurality of HPT stator vanes; 
 wherein a core gas path extends through the HPC, the combustor section, the HPT, and exits through the HPC discharge; 
 wherein a gas leakage flow is a bled gas off of the core gas path forward of the HPC discharge; 
 a bypass tangential on board injector (TOBI) system that extends circumferentially around the engine axial centerline, the bypass TOBI system having a plurality of nozzles, an inner radial side, an outer radial side, a plurality of first radial passages, and a plurality of second radial passages, 
 wherein the gas turbine is configured such that a portion of core gas exiting the core gas path through the HPC discharge is directed to the bypass TOBI system as a diffuser inner diameter (ID) flow that passes through the plurality of nozzles, 
 wherein the configuration of the plurality of first radial passages is different from the configuration of the plurality of second radial passages where the second radial passages each have a fluid flow path that directs the bled gas in an upstream direction with respect to the ID flow and the first radial passages do not direct the bled gas in the upstream direction, 
 wherein the plurality of first radial passages are configured to allow the bled gas from the HPC to pass from the inner radial side of the bypass TOBI system to the outer radial side of the bypass TOBI system without mixing with the diffuser ID flow passing through the plurality of nozzles, 
 and wherein the plurality of second radial passages are configured to allow the bled gas from the HPC to pass from the inner radial side of the bypass TOBI system to the outer radial side of the bypass TOBI system without mixing with the diffuser ID flow passing through the plurality of nozzles. 
 
     
     
       2. The gas turbine engine of  claim 1 , further comprising a first gas port in communication with the HPC, the first gas port configured to receive the bled gas, and a first HPC gas passage that provides fluid communication between the first gas port and the bypass TOBI system. 
     
     
       3. The gas turbine engine of  claim 2 , wherein the HPC has “N” number of HPC rotor stages, wherein “N” is an integer equal to or greater than two, and the “N” number of HPC rotor stages includes a first HPC rotor stage disposed closest to an axial forward end of the HPC and an Nth HPC rotor stage that is the HPC rotor stage closest to the HPC discharge;
 wherein the first gas port is disposed in the HPC aft of the Nth HPC rotor stage. 
 
     
     
       4. The gas turbine engine of  claim 3 , wherein the first gas port and the first HPC gas passage are configured such that substantially all of the bled gas received by the first gas port is directed to pass from the inner radial side of the bypass TOBI system to the outer radial side of the bypass TOBI system. 
     
     
       5. The gas turbine engine of  claim 4 , wherein the HPT stator vane stage includes a first HPT stator vane stage disposed adjacent the combustor section and the HPT rotor stage is a first HPT rotor stage disposed aft of the first HPT stator vane stage, and a FS-HPT rotor forward rim cavity is disposed aft of the first HPT stator vane stage, forward of the first HPT rotor stage, and radially inside of the core gas path; and
 the gas turbine engine is configured such that the bled gas is directed into the FS-HPT rotor forward rim cavity for exit into the core gas path forward of the first HPT rotor stage after passing radially through the bypass TOBI system. 
 
     
     
       6. The gas turbine engine of  claim 5 , wherein the combustor section includes a diffuser outer diameter (OD) flow path disposed radially outside of the combustor, and wherein the diffuser ID flow passes though a diffuser inner diameter (ID) flow path disposed radially inside of the combustor. 
     
     
       7. The gas turbine engine of  claim 6 , wherein the combustor section is configured such that a first portion of gas exiting the HPC discharge is directed into the diffuser OD flow path; and
 the engine is configured to direct the first portion of gas outside of the core gas path to the HPT. 
 
     
     
       8. The gas turbine engine of  claim 7 , wherein the HPT further includes a second HPT stator vane stage disposed aft of the first HPT rotor stage; and
 at least a portion of the first portion of gas exiting the HPC discharge is directed to the second HPT stator vane stage. 
 
     
     
       9. The gas turbine engine of  claim 6 , wherein the diffuser ID flow path is in fluid communication with the plurality of nozzles of the bypass TOBI system to direct the diffuser ID flow through the plurality of nozzles;
 wherein the engine is configured to direct the diffuser ID flow to the first HPT rotor stage. 
 
     
     
       10. The gas turbine engine of  claim 9 , wherein the plurality of first radial passages are circumferentially spaced apart from one another around the bypass TOBI system, and the plurality of second radial passages are circumferentially spaced apart from one another around the bypass TOBI system. 
     
     
       11. The gas turbine engine of  claim 10 , wherein the plurality of first radial passages are disposed at a forward end of the bypass TOBI system and radially through the bypass TOBI system to a first outer radial compartment. 
     
     
       12. The gas turbine engine of  claim 11 , wherein a plurality of seals are disposed on the inner radial side of the bypass TOBI system and are engaged with an inner radial surface of the bypass TOBI nozzles, wherein a first seal of the plurality of seals and a second seal of the plurality of seals are spaced apart from one another and define an annular region therebetween;
 wherein said second radial passages extend from the annual region and radially through the bypass TOBI system to a second outer radial compartment. 
 
     
     
       13. The gas turbine engine of  claim 2 , wherein the HPC has “N” number of HPC rotor stages, wherein “N” is an integer equal to or greater than two, and the “N” number of said HPC rotor stages includes a first HPC rotor stage disposed closest to an axial forward end of the HPC, an Nth HPC rotor stage that is the HPC rotor stage closest to the HPC discharge, and an N−1th HPC rotor stage disposed between the first HPC rotor stage and the Nth HPC rotor stage; and
 a second gas port in communication with the HPC, the second gas port configured to receive gas from the core gas path forward of the N−1th HPC rotor stage; and 
 wherein the gas received from the second gas port (HPC bleed gas) is routed through a passage disposed radially outside of the core gas path to the HPT. 
 
     
     
       14. The gas turbine engine of  claim 13 , wherein the second gas port is in communication with an outer wall segment forward of the N−1th HPC rotor stage; and
 the HPC bleed gas received from the second gas port is routed to the HPT in unmixed form. 
 
     
     
       15. A gas turbine engine having an axial centerline, comprising:
 a high pressure compressor (HPC); 
 a combustor section having a combustor; 
 a high pressure turbine (HPT); 
 wherein a core gas path extends through the HPC, the combustor section, the HPT, and exits through an HPC discharge; 
 wherein a gas leakage flow is a bled gas off of the core gas path forward of the HPC discharge; 
 a bypass tangential on board injector (TOBI) system that extends circumferentially around the engine axial centerline, the bypass TOBI system having a plurality of nozzles, an inner radial side, an outer radial side, a plurality of first radial passages, and a plurality of second radial passages, 
 wherein the gas turbine is configured such that a portion of core gas exiting the core gas path through the HPC discharge is directed to the bypass TOBI system as a diffuser inner diameter (ID) flow that passes through the plurality of nozzles, 
 wherein the configuration of the plurality of first radial passages is different from the configuration of the plurality of second radial passages where the second radial passages each have a fluid flow path that directs the bled gas in an upstream direction with respect to the ID flow and the first radial passages do not direct the bled gas in the upstream direction, 
 wherein the plurality of first radial passages are configured to allow the bled gas from the HPC to pass from the inner radial side of the bypass TOBI system to the outer radial side of the bypass TOBI system without mixing with the diffuser ID flow passing through the plurality of nozzles, 
 and wherein the plurality of second radial passages are configured to allow the bled gas from the HPC to pass from the inner radial side of the bypass TOBI system to the outer radial side of the bypass TOBI system without mixing with the diffuser ID flow passing through the plurality of nozzles. 
 
     
     
       16. The gas turbine engine of  claim 15 , further comprising a first gas port in communication with the HPC, the first gas port configured to receive the bled gas, and an HPC gas passage that provides fluid communication between the first gas port and the bypass TOBI system;
 wherein the HPC has “N” number of said HPC rotor stages, wherein “N” is an integer equal to or greater than two, and the “N” number of said HPC rotor stages includes a first HPC rotor stage disposed closest to an axial forward end of the HPC and an Nth HPC rotor stage that is the HPC rotor stage closest to the HPC discharge; 
 wherein the first gas port is disposed in the HPC aft of the Nth HPC rotor stage; and 
 wherein the first gas port and the HPC gas passage are configured such that substantially all of the bled gas received by the first gas port is directed to pass from the inner radial side of the bypass TOBI system to the outer radial side of the bypass TOBI system. 
 
     
     
       17. The gas turbine engine of  claim 16 , wherein a plurality of HPT stator vane stages includes a first HPT stator vane stage disposed adjacent the combustor section and a plurality of HPT rotor stages includes a first HPT rotor stage disposed aft of the first HPT stator vane stage, and a FS-HPT rotor forward rim cavity is disposed aft of the first HPT stator vane stage, forward of the first HPT rotor stage, and radially inside of the core gas path; and
 the gas turbine engine is configured such that the bled gas is directed into the FS-HPT rotor forward rim cavity for exit into the core gas path forward of the first HPT rotor stage after passing radially through the bypass TOBI system. 
 
     
     
       18. The gas turbine engine of  claim 15 , wherein each of the first radial passages includes a first seal leak path portion disposed on the outer radial side of the bypass TOBI system. 
     
     
       19. The gas turbine engine of  claim 18 , wherein each of the second radial passages includes a second seal leak path portion disposed on the inner radial side of the bypass TOBI system. 
     
     
       20. A method for operating the gas turbine engine of  claim 1 , the method comprising:
 receiving gas from the HPC; 
 passing the HPC gas radially through the bypass TOBI system from the inner radial side to the outer radial side, wherein the HPC gas passes through the bypass TOBI system from the inner radial side to the outer radial side by way of the plurality of first radial passages and the plurality of second radial passages, the plurality of first radial passages and the plurality of second radial passages independent of one another; and 
 purging a cavity disposed between a first the HPT stator vane stage and a first the HPT rotor stage with the HPC gas, the HPC gas entering the core gas path forward of the first HPT rotor stage thereafter.